E.CUBE MISSION: THE ENVIRONMENTAL CUBESAT

Camilla Colombo(1), Mirko Trisolini(1), Francesca Scala(1), Marco Paolo Brenna(1), Juan Luis Gonzalo(1), Stefano Antonetti(2), Francesco Di Tolle(2), Roberto Redaelli(2), Francesco Lisi(2), Luca Marrocchi(3), Marco Alberti(3), Alessandro Francesconi(4), Lorenzo Olivieri(4), Massimo Tipaldi(5)

(1) Dep. of Aerospace Science and Technology, Politecnico di Milano, Milano, Italy, Email: [email protected] (2) D-Orbit SpA, Viale Risorgimento 57, 22073 Fino Mornasco, Como, Italy, Email: [email protected] (3) Temis, Via G. Donizetti, 20, 20011 Corbetta (MI), Italy, Email: [email protected] (4) University of Padova, Padova, Italy, Email: [email protected] (5) Intelligentia, Via Del Pomerio 7, 82100 Benevento, Italy, Email: [email protected]

collected data about non-trackable fragment objects, (3) ABSTRACT to characterise the atmosphere for more accurate re-entry The e.Cube mission aims at contributing to the predictions and the thermomechanical loads experienced advancement of technologies and methodologies by the spacecraft during re-entry. dedicated to space debris mitigation and remediation (1) A 12U CubeSat will be deployed in Low Earth Orbit to increase spacecraft autonomy in performing CAMs, (LEO) where the operational phase will be dedicated to (2) to support space debris modelling with in-orbit three experiments. During the first phase, the collected data about non-trackable fragment objects, (3) autonomous CAM experiment will be carried out, to characterise the atmosphere for more accurate re-entry consisting of several in-flight CAM tests for simulated predictions and the thermomechanical loads experienced close approaches with a virtual debris. For each test, a by the spacecraft during re-entry. This paper presents the sequence of virtual Conjunction Data Messages (CDM) preliminary mission analysis, payload selection and will be generated on ground and uploaded to the system design for the mission. spacecraft. The CAM payload processor will then decide if and how to perform the CAM making use of artificial INTRODUCTION intelligence. A particle detection device, exposed in the The space surrounding our planet is densely populated by frontal side of the spacecraft, will collect over the time of an increasing number of space debris, whose number is one year the highest amount possible of sub-millimetre expected to grow in the next decade due to in-orbit particles to be used to validate in-use statistical explosions, material deterioration, and in-orbit collisions. distributions of non-trackable objects. Finally, during the Space debris poses a threat to the current and future last phase of the mission an EOL manoeuvre will be access to space. While debris objects larger than 5-10 cm implemented. During the re-entry phase an experiment can be tracked from Earth and need to be avoided through through the re-entry data collector will characterise the the execution of Collision Avoidance Manoeuvres thermosphere in the region below 200 km, through some (CAMs), smaller debris are not currently catalogued, in-situ measurements, in particular temperature and therefore statistical distributions are used to quantify the pressure. collision risk in different orbital regions and to define The role of the e.Cube mission is to contribute to the mitigation guidelines through long-term evolution advancement of technologies and methodologies models. Moreover, while the design of End-Of-Life dedicated to space debris mitigation and remediation. (EOL) disposal manoeuvres is desirable to mitigate the Three key areas have been identified: the development of number of inoperative satellites, uncertainties related to autonomous CAM algorithms, the detection and our limited knowledge of the solar activity and its measurement of untraceable space debris, and the interaction with the atmosphere make accurate re-entry characterisation of the upper atmosphere and of the predictions a challenging task [1]. satellite disposal mechanism. These three emerging areas The e.Cube mission [2] aims at contributing to the are also the research pillars of the COMPASS project, advancement of technologies and methodologies funded by the European research Council at Politecnico dedicated to space debris mitigation and remediation (1) di Milano [3]. The tools and techniques developed within to increase spacecraft autonomy in performing CAMs, this project have been already brought to their proof of (2) to support space debris modelling with in-orbit concept thanks to projects funded by the European Space

LProc.eave 8th foo Europeanter em pConferencety – The on C onSpacefer eDebrisnce f (virtual),ooter w Darmstadt,ill be add Germany,ed to th 20–23e firs tApril page 2021, of epublishedach pa pbye rthe. ESA Space Debris Office Ed. T. Flohrer, S. Lemmens & F. Schmitz, (http://conference.sdo.esoc.esa.int, May 2021) Agency (ESA) and the Italian Space Agency (ASI), for Space Safety (MISS) software tool which couple the however the status of the such algorithms is still at the uncertainties in the environment model and orbit preliminary operational software development. It is the determination measurements [4]. Supervised artificial aim of the e.Cube mission to develop operational intelligence techniques will be employed for the planning software for Space flight and on-orbit experiments, and and decision making in the operations for autonomous to send back to Earth relevant scientific data that will be Collision Avoidance Manoeuvres (CAM). During the used to validate long-term evolution models for space CAM in-flight experiment, synthetic conjunction data debris, re-entry prediction, and to improve autonomous messages, simulating the possible collision threat with a CAM for the future Space Traffic Management (STM). debris, will be transmitted to the spacecraft and the CAM The data and results of the e.Cube mission will be used command module will autonomously decide, compute, to validate tools and techniques currently used within the and command the required manoeuvre to be performed space debris community: long-term evolution models for [5]. An expected output of this objective is to advance the space debris, re-entry prediction, and to improve proposed CAM algorithms to autonomy level E4 E4 autonomous CAM for the future space traffic (goal commanding) according to “ECSS-E-ST-70-11C – management. In line with the ideal of the e.Cube mission Space segment operability”. to contribute to the international discussion and cooperation on space debris mitigation, all the data and Obj. 2 – DEBRIS. Characterisation of results of the on-board experiments will be freely shared untraceable space debris objects to update and with the whole debris community to be used for improve space debris environmental models advancing the definition of space debris mitigation (BRIS). guidelines. Obj. 2 will be fulfilled through the design of a particle detection device, which will characterise in-situ sub- MISSION OBJECTIVES millimetre level particles in Low Earth Orbit by sensing The e.Cube mission aims at participating to the effort in the frequency, mass, energy, and direction of debris Space Situational Awareness and Space Traffic particles. The Particle Detection Device (PDD) will be Management, by providing a significant contribution to pointed in the velocity direction with 1-degree accuracy, the development of key areas with the aim of ensuring a to maximise the particle collection. To properly safer and more sustainable access to Space. In the rapidly characterise the sub-millimetre debris (less than 1 mm) evolving scenario of New Space activities, large and meteoroids environment, the PDD shall constellations, and the increasing dependence of our collect at least 200 particles during the spacecraft daily life on Space, space agencies have identified the operating life. mitigation of space debris as one of the main goals for a sustainable development of space activities in the near Obj. 3 – RE-ENTRY. Characterisation of the future. The scientific objectives of the e.Cube mission, upper atmosphere for more accurate re-entry represented in Figure 1, address distinct but interlinked prediction and of the thermomechanical loads aspects of space sustainability, spanning the entire experienced by the spacecraft during re-entry. lifetime of any space mission. The first two objectives Obj. 3 aims at reducing the model uncertainties on the investigate the two sides of the same spectrum: the post-mission disposal, and specifically on the re-entry collision avoidance of traceable debris on one side and phase, which arise due to the atmospheric modelling the improvement of the damage assessment from (especially the solar activity) and to the satellite demise untraceable debris on the other. They will be explained in and break-up process, a process for which little to no the next Sections. mission-related data is currently available. This objective will be achieved by arranging a distributed network of Obj. 1 – CAM. Development, validation and sensors inside the spacecraft that will measure the testing of on-board algorithms for autonomous evolution of the acceleration, pressure, and temperature collision avoidance during the re-entry phase to record the mechanical and To demonstrate autonomous collision avoidance thermal loads suffered by the spacecraft. The experiment capabilities during at least three relevant in-flight shall collect data in the region between 200 km and 100 scenarios. This will be achieved by implementing the km of altitude with a resolution of at least 10 km. The semi-analytical collision avoidance algorithms CubeSat shall transmit at least 125 kbit/s data to the developed at PoliMi within the Manoeuvre Intelligence ground until an altitude of 100 km.

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Figure 1. Conceptual representation of the e.Cube objectives.

regulations [7, 8], that require a decay of the satellite MISSION ANALYSIS within 25 years (see Figure 3). For this reason, the The selection of the mission and spacecraft architecture, operational orbit will be a SSO below 550 km of altitude the operational orbit, and the disposal strategy are to be compliant with the space debris mitigation fundamental aspects for the e.Cube mission. The selected regulations even without disposal manoeuvre, which is a configuration is a 12U CubeSat, which can best strong requirement, given the objectives and background accommodate the three payloads required for the of the mission. fulfilment of the objectives. Given the distribution of space debris in the LEO region, the selection of the operational orbit directly influences Obj. 2. In fact, different orbital regions are characterised by different debris fluxes, which have a direct impact on the number of particles that can be analysed by the payload. Figure 2 shows the value of debris fluxes as a function of semimajor axis and inclination as computed with ESA MASTER 8 [6]. The regions with the highest fluxes correspond to Sun-Synchronous Orbits (SSO).

Figure 3. Natural decay time vs starting altitudes. Circular SSO orbits and solar flux uncertainties.

Table 1 summarises the estimated number of particles impacts for a one-year mission for orbits between 400 km and 800 km with a 12U CubeSat (cross-section of 0.2 x 0.2 m2). Particles with a diameter between 1 μm and 1 cm are considered in Table 1 as they are untraceable by current ground observation facilities [1]. Although a 500 km altitude orbit can detect less than half the total Figure 2. Debris fluxes for orbits between 400 km and impacts with respect to an 800 km orbit, this amount is in 1000 km. line with the proposed target of 200 particles, while providing a safe EOL disposal considering a natural These orbits are the perfect candidates for a debris decay. Therefore, it is proposed an operational orbit of characterisation mission as they provide the possibility to SSO type between 500 km and 550 km of altitude and a collect a significant amount of data. Additionally, they mission lifetime of at least 1 year, to reach the target of have directional fluxes (concentrated in the front part of detected particles. the spacecraft), which allow for a better detection of As Obj. 3 aims at characterising the re-entry of the particles and design of the payload. Even if peak-flux can spacecraft, it is important to design a proper disposal be experienced for orbits in the range 800-900 km, they scenario. As the disposal phase is also part of the would not be compliant, in case of failure of the on-board operational phase of the mission, it must be carried out propulsion system, with the space debris mitigation within a timeframe that is compatible with the lifetime

Leave footer empty – The Conference footer will be added to the first page of each paper. and reliability of CubeSat components. Therefore, a The apogee altitude is the starting altitude of the maximum disposal time of 6 months has been operation orbit and the perigee is the target for the Re- considered. The CubeSat will perform disposal entry Analysis Phase. It was observed that starting from manoeuvres to lower its perigee altitude and ensure the a 550 km orbit, with about 100 m/s delta-v it is possible compliance with this requirement. to reach a perigee of about 180 km. During Phase A, a more detailed disposal analysis will be performed that Table 1. Number of impacts per year on different will also consider a multi-burn strategy, to better operational orbits of SSO type. distribute the delta-V of each burn and reduce the effects of possible misalignments or failures. Altitude Meteoroid Debris Total [km] impacts impacts Impacts [#/year] [#/year] [#/year] TECHNOLOGYCAL READINESS OF THE MISSION 400 ~ 122 ~ 10 ~ 132 500 ~ 125 ~ 70 ~ 195 Deriving from the objectives of the mission defined in the 600 ~ 126 ~ 180 ~ 306 previous Section, the key technologies for the mission 800 ~ 128 ~ 320 ~ 448 success consist of the three payloads dedicated to the fulfilment of the objectives, and of some subsystems whose technologies are qualifying for the success of the Another important aspect in view of the fulfilment of Obj mission. Table 2 summarises the requirements of the 3 is the time spent in the lower parts of the atmosphere to main systems and subsystems of the CubeSat. maximise the acquisition of relevant data during the re- entry. The delta-V required for the disposal from the operational orbit with a single burn was computed [2].

Table 2. Summary of main system and subsystem requirements.

ID Requirement CAM-1 The payload shall compute autonomously on-board the optimal CAM given the collision alert message (i.e. impactor state, covariance, and warning time). CAM-2 The CAM shall be performed in less than one quarter of the period from the predicted close approach. DEB-1 The PDD shall be able to detect sub-millimetre particles between 1 μm and 1 mm and characterise their dimension, velocity, and direction. DEB-2 The PDD shall detect at least 200 particles, including both debris and meteoroids, during the operational time. RDC-1 The RDC shall be able to measure the pressure and temperature of the atmosphere between 200 km and 100 km with a resolution of 10 km. RDC-2 The RDC shall be able to characterise the thermal and mechanical loads sustained by the CubeSat during the disposal phase. SYS-1 The CubeSat mass shall be less than 24 kg. SYS-2 The CubeSat volume shall be less or equal than 12U. AOCS-1 The AOCS shall provide a pointing accuracy of 1 deg during the Debris Analysis Phase. AOCS-2 The AOCS shall provide a maximum slew rate of 1 deg/s. AOCS-3 The AOCS shall provide a pointing accuracy of 5 deg during the Re-entry Analysis Phase. PROP-1 The propulsion system shall provide at least 125 m/s for CAM and Re-entry Phase. PROP-2 The maximum thrust of the propulsion system shall be 500 mN. EPS-1 The EPS shall supply an average of 22 W and a peak of 29 W during the Debris Analysis Phase. EPS-2 The EPS shall supply an average of 17 W and a peak of 53 W during the CAM phase. EPS-3 The EPS shall supply an average of 18 W and a peak of 25 W during the Re-entry Analysis Phase. COMM-1 The CubeSat shall be able to send telemetry data in the UHF/VHF frequency band for space application. COMM-2 The communication system shall be able to download at least 600 kB/day during the Debris Analysis Phase, and at least TBD kB/orbit during the Re-entry Analysis Phase.

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commands do not exceed a predefined operational E.CUBE MISSION PAYLOADS envelope. Reference [5] shows a schematic In this Section an overview of the payload technologies, representation of the CCM and the CAM software. The their preliminary design, and their TRL will be given. To hardware selected for the CCM is D-Orbit’s OBC Core, increase the modularity and reliability of the payloads, which is also selected as the OBC in the OBDH system. each one of them will carry a dedicated on-board OBC Core has a TRL of 9, and it has been successfully computer, which will process the acquired data. Table 3 implemented on the ION spacecraft. The CAM algorithm summarises the current TRL level of the payloads and the is currently at TRL 4, and during phase A/B it will be aimed TRL for Phase A/B. brought up to TRL 6 before its implementation in the CCM. Regarding the autonomy level, the objective is to Table 3. TRL level of the payloads. demonstrate “ECSS-E-ST-70-11C – Space segment operability” level E4 autonomy during the in-orbit CAM Payload Current TRL Target TRL experiment. for phase A/B CAM HW 9 / SW 4 HW 9 / SW 6 A preliminary estimation of the design parameters of the Command CCM is given in Table 4. The size, mass and power are Module based on those of the OBC Core. The expected data Particle 3/4 5 budget it very limited, including only the CDMs Detection (maximum 3-4 per day) in uplink and the log of the CCM Device operation in downlink. Re-entry Data 3/4 5 Table 4: Preliminary design of the CCM payload. Collector Parameter Value CAM Command module Size 15 x 15 x 2.5 cm3 The payload for the CAM experiment is the CCM, a Mass < 500g dedicated OBC implementing the algorithms for on- Power 0.4 W (maximum 1.8 W) board CAM decision-making and design. The CCM only Data volume 20 kB/day interacts directly with the CubeSat’s OBC, receiving the synthetic CDMs sent from ground and navigation Particle Detection Device (PDD) information, and generating a manoeuvre command In-situ space debris and meteoroids detectors developed when it determines a CAM is required. The CAM so far are based on several physical principles. Among command will be handled by the OBC like a command those, the most common systems employ strip-film sent from ground to perform a manoeuvre. The CCM will conductive and piezoelectric materials. On one hand, also log the key parameters related to the CAM decision- strip-film sensors consist in a thin film of non-conductive making and design process, and they will be transmitted material on which thin conductive stripes are deposed in to ground for analysis. The main objective of the CAM parallel; in case of impact, the film is perforated, and one experiment is to advance the TRL and autonomy level of or more stripes are cut and their electrical continuity is the on-board CAM algorithms and prove they can be used interrupted [9]. By employing two layers it is possible to in future operational satellites. To this end, two criteria detect the position of the impact event, while the extent will be considered during the design of the CCM: 1) to of the damage of stripes is used to estimate the properties make it as independent as possible from the platform, and of the impactor, and the sensitivity of the sensor depends 2) to clearly define the data interfaces with the OBC. The on the density of stripes. On the other hand, piezoelectric on-board CAM software is composed of 2 modules: the detectors employ patches of piezoelectric materials (e.g. decision-making module and the CAM design module. PVDF [10]) and measure the electric spike due to charge The decision-making module determines whether a CAM separation following the material deformation during the is needed or not to keep the collision probability at close impact transient. The detection capability of piezoelectric approach below a given threshold (typically 10-4) based sensors is currently in the range of 1 µm to 1 mm at a on the sequence of CDMs and navigation information. It velocity of up to 10 km/s ; its sensitivity can be set by the is based in machine learning algorithms, trained on conditioning electronics. Other systems proposed for ground with historical and synthetic CAM datasets. The debris impact detection come from direct heritage of CAM design module implements semi-analytical models scientific instruments for dust collection used in several for the efficient computation of maximum deviation or space missions (e.g. Cassini, Stardust, BepiColombo, a minimum collision probability impulsive CAMs2. The detailed list is reported in Bauer et al. [11]) and employ software implemented in the CCM will also include a acoustic sensors or piezoelectric panels to measure the supervision module, tasked with: 1) logging the principal momentum transferred to the payload during impact parameters related to the operation of the other two events. Such systems were developed to detect mostly modules, and 2) verifying that the generated CAM dust and micrometeoroids with impact velocities and

Leave footer empty – The Conference footer will be added to the first page of each paper. fluxes larger than those typical of space debris and Power 5 W require complex conditioning electronics. A sketch of the Data volume 200 kB/day PDD that we are going to develop within e.Cube is shown in Figure 4. The PDD design will use existing solutions Re-entry Data Collector (RDC) as baseline. During the execution of the e.Cube, the PDD technology maturity will be continuously increased The RDC has two main objectives: through breadboarding and calibration activities, until the • The measurement of the density, below 200 km realisation and testing of a payload’s proto-flight model altitude, during Phases C/D. • The measurement of thermal and mechanical loads in the final mission phase before the re-entry. First, the knowledge of the air density in the thermosphere is crucial for improving the accuracy of the SW models used to estimate the re-entry trajectory of debris. A first approach is identified in the direct measurement of the density in situ, which was typically performed with expensive payloads, not suitable for a CubeSat mission. Nevertheless, a second indirect approach can be considered: the density can be computed also as an indirect parameter after the measurement of other quantities, in particular temperature and pressure. Figure 4: PDD schematic. The latter approach is suitable for CubeSat mission, being the temperature and pressure sensors for CubeSat already As baseline, we will evaluate the use of a piezoelectric available on the market, and they do not need any detector sandwiched with two layers of conductive development or customisations. The second objective of stripes. This combination will make possible to both the RDC is the understanding of the thermal and cross-validate the two sensors and extend the working mechanical loads acting on the CubeSat during the latest range of the detector through a proper adjustment of the phase of its life. Inertial Measurement Units (IMUs) sensitivity of the two elements. The detector will be based on Micro Electro-Mechanical Systems (MEMS) backed by a soft material plate employed as debris technology are the state-of-the-art and several catcher to avoid secondary signals due to debris cloud Commercial off-the-shelf (COTS) options are currently reflections and mounted on the satellite external available on the market. The inclusion of temperature structure. A first sensor breadboard will be developed in sensors and strain gauges to measure the internal heat the flat-sat format for design-optimization purposes and loads and displacements, respectively, will be considered will feature a scaled prototype of the PDD plate to be in Phase A, when a more consolidated internal design of subjected to electric and functional testing. The driving the CubeSat will be available. and conditioning electronics will be realized as well. A second breadboard will consist in a more compact Since TEMIS heritage on space payload is based on electronic box and a series of instrumented PDD plates systems larger than CubeSats, the currently developed with different configurations (e.g. thickness, layers products cannot be directly implemented in the CubeSat, material), to be evaluated by hypervelocity testing for full but they require some customisations. For the RDC, a characterisation. The impact test campaign will make it dedicated custom CPU board will be realised, exploiting possible to point out the best design solution and will the design heritage, including the COTS sensors and calibrate the sensor in different working conditions in drivers to reduce costs and risks. The following sensors terms of impact momentum, debris size and velocity. It are considered in this preliminary assessment: is underlined that calibration results could be also • Pressure sensor: A set of three pressure sensors will employed to improve the analysis of data collected by be used. As baseline, the MKS 905 MicroPirani other sensors already in space. A preliminary estimate of sensors have been selected [2]. the PDD SWAP in given in the Table 5: • Heat Flux Sensor: From the heat flux in the velocity Table 5: Preliminary design of the PDD payload. direction the information about temperature of the Parameter Value flux can be computed. For the baseline, the Omega Size (PDD plate) 18 x 18 x 2 cm3 HFS-3 heat flux sensor has been selected [2]. Size (PDD electronics) 9.2 x 9.7 x 4 cm3 (CubeSat standard) PRELIMINARY SYSTEM DESIGN Total mass 1.3 kg The following sections present the preliminary mission (plate+electronics) design of the CubeSat for the main subsystems,

Leave footer empty – The Conference footer will be added to the first page of each paper. considering the requirements arising from the payloads following actuators architecture is considered in this and the operational phases. When applicable, possible preliminary assessment: options for the equipment selection have been given, • alongside the TRL of the components. Given the large Reaction Wheels: the RWs are selected among the availability of COTS with low prices and space heritage, commercial flight-proven available on the market. the approach here is to select systems and components Four RWs, one for each axis plus a redundant one, are from Italian and European manufacturers. This reduces selected to make a 3-axis attitude control [2]. risks and allows to focus on the scientific and • Magnetorquers: the magnetorquers are selected technological nature of the mission. among the commercial flight-proven available on the market. A three-axis magnetorquer will be carried on- Structure and configuration board. They will be used to perform desaturation of The structure of the spacecraft is a standard 12U the RWs [2]. CubeSat, which is compatible with the integration of • Attitude control module: the possibility of including COTS components. The preliminary configuration of the a compact cold-gas thruster module will be evaluated e.Cube spacecraft is shown in Figure 5. Some highlights during the Phase A of the mission if needed during are the presence of the PDD on the front face, and two the desaturation of the RWs. These modules can deployable solar panels to guarantee the power provide up to 10 mN of thrust along 2-axis at low production. On the back, the two main engines used for power consumption (< 2W). CAM and disposal manoeuvre are visible. The attitude determination will be performed with flight- proven sensors. Given the currently available sensors for CubeSats, the expected accuracy of the attitude determination system is up to 0.05 degrees when star- trackers are included. A summary of the main components for attitude determination is given in [2]. The selection of the actuators is driven by the requirements related to the three operational phases, shown in Table 2. Preliminary analyses have been performed to estimate the maximum momentum storage and the maximum torque required during each one of the phases.

Propulsion system Figure 5. Front view of the preliminary configuration of The propulsion system design is crucial to the mission the e.Cube spacecraft. success as it will perform both the CAM and the disposal manoeuvres before the re-entry analysis phase. Given the Attitude and orbit control system requirements PROP-1 and PROP-2 (Table 2) a trade-off analysis has been performed on currently available The Attitude and Orbit Control System (AOCS) is a CubeSat propulsion systems. The CAM experiment crucial subsystem for the success of the mission. During requires a minimum delta-V of 5 m/s. Considering a 5- the CAM experiment the system shall ensure the correct minute burn to perform the CAM, the maximum required orientation of the CubeSat to perform the manoeuvre in thrust level for one manoeuvre is 80 mN. The maximum the direction specified by the algorithm. During the delta-V required to change the orbit for the Re-entry debris detection experiment, the ACDS shall ensure the Analysis Phase is 120 m/s, including margins. RAM pointing of the CubeSat with an accuracy of 1 degree. During the re-entry analysis experiment, the Possible candidates for the propulsion system have been system shall be able to maintain a data relay connection identified [2]. Given the delta-V requirements for the to download the acquired data until an altitude of 100 km. optimal strategy for the Re-entry Analysis Phase, the To ensure a correct attitude of the CubeSat during the baseline propulsion system design foresees a set of two mission lifetime, the system is controlled on the three chemical thrusters. This option also guarantees a axes with Reaction Wheels (RW). The de-tumbling is redundancy of the propulsion system and a sub-optimal performed with magnetorquers. The desaturation control disposal strategy for the Re-entry Analysis Phase in case of the RWs is performed using together magnetorquers of failure of one of the thrusters. In addition, it also allows and, if needed, CubeSat attitude thruster module, relying a more responsive system during the CAM phase. An on cold-gas thrusters, aligned with the Centre of Ma, thus electric option can also be considered and in this case a minimising parasitic torques (nominally null). The

Leave footer empty – The Conference footer will be added to the first page of each paper. single thruster will be used. A more detailed trade-off A preliminary sizing of the solar arrays was performed analysis will be performed during the Phase A. [2]. Considering the 22 W average power case, the required array area is about 900 cm2. The analysis was Electrical power system performed considering as baseline orbit a dawn-dusk The electrical power system architecture for the e.Cube SSO to minimize the eclipse load on solar arrays and on- mission is based on solar arrays, secondary batteries and board batteries. During Phase A, a trade-off will be the power conditioning and distribution electronics. The performed to analyse the possibility to carry the mission EPS shall be sized based on the power and energy on a different orbital configuration. During the Debris required during each mission phase. Power requirements Analysis Phase and the Re-entry Analysis Phase, the for each mission phase are summarized in Table 6. CubeSat will be RAM pointing; therefore, the solar panels will be placed on the 2-by-3 lateral side facing the Table 6: Power budget summary for the main operational Sun, together with a solar panel wing, to grant the desired phases. level of power for the system. Additionally, to ensure the Phases Avg. Power Peak Power power generation in case of pointing misalignment and (W) (W) for redundancy reasons, solar cells will be placed on CAM Phase 17 53 every lateral side of the CubeSat. Debris Analysis 22 29 Phase Mass budget Re-entry 18 25 A preliminary Mass Budget is presented in Table 7, Analysis Phase considering possible for the system components from the analysis presented above.

Table 7: Preliminary mass budget.

SubSys Components Mass (g) Margin Mass with margin (g) P/L PDD Payload 1300 20% 1560 CAM Payload 500 20% 600 RDC Payload 500 20% 600 STR Structure 1750 20% 2100 PS Propulsive System 2820 20% 3384 AOCS Sun Sensors 8.8 20% 10.56 Star Tracker 106 20% 127.2 Gyroscope 55 20% 66 Magnetometer 8 20% 9.6 Magnetorquer 106 20% 127.2 Reaction Wheels 940 20% 1128 GNSS receiver 20 20% 24 GNSS antenna 18 20% 21.6 C&DH Main Computer 150 20% 180 TT&C UHF Antenna 115 20% 138 S-Band Antenna 110 20% 132 UHF Transceiver 94 20% 112.8 S-band Transceiver 270 20% 324 EPS PCDU 191 20% 229.2 TCS Thermal 315 20% 378 Harness 315 20% 378 Total 9061.8 20% 10874.16

to increase spacecraft autonomy in performing CAMs, CONCLUSION (2) to support space debris modelling with in-orbit The preliminary mission design for the e.Cube mission is collected data about non-trackable fragment objects, (3) presented. The e.Cube mission aims at contributing to the to characterise the atmosphere for more accurate re-entry advancement of technologies and methodologies predictions and the thermomechanical loads experienced dedicated to space debris mitigation and remediation (1) by the spacecraft during re-entry. The proposing team

Leave footer empty – The Conference footer will be added to the first page of each paper. will engage with other European and extra European Fragmentation and Environment Evolution Models,” companies, institution and agencies for enlarging the Final Report, Contract N. 4000115973/15/D/SR, e.Cube consortium. Institute of Space System, Technische Universität Braunschweig, 26/08/2020 REFERENCES 7. European Space Agency, “Requirements on Space 1. Colombo C., Letizia F., Trisolini M., Lewis H. “Space Debris Mitigation for ESA Projects”, 2008 Debris: Risk and Mitigation”, In Frontiers of Space 8. European Space Agency, “Space Debris Mitigation Risk, pp. 105-141, CRC Press, 2018 Compliance Verification Guidelines”, 2015. 2. Colombo C., Trisolini M., Scala F., Brenna M. P., 9. Nakamura M., et al., "Development of in-situ micro- Gonzalo J. L., Antonetti S., Di Tolle F., Redaelli R., debris measurement system," Advances in Space Lisi F., Marrocchi L., Alberti M., Francesconi A., Research, Vol. 56, Issue 3, pp. 436-448, 2015. Olivieri L., Tipaldi M., “e.Cube – The environmental CubeSat”, Proposal to the Italian Space Agency, 10. Laufer R., et al. "ARMADILLO - A 29/11/2020 Demonstration for Low-Cost In-Situ Investigations of the Upper Atmosphere of Planetary Bodies," 3. COMPASS ERC, www.compass.polimi.it, Last Proceedings of International Planetary Probe accessed 08/04/2021. Workshop (IPPW). 4. Gonzalo J.L., Colombo C., Di Lizia P., “Introducing 11. Bauer W., et al. "Development of in-situ space MISS, a new tool for collision avoidance analysis and debris detector," Advances in Space Research, Vol. 54 design”, Journal of Space Safety Engineering, Vol 7, Issue 9, pp. 1858-1869, 2014 pp. 282-289, 2020. https://doi.org/10.1016/j.jsse.2020.07.010 ACKNOWLEDGEMENTS 5. Gonzalo J.L., Colombo C., “On-Board Collision The research to propose the e.Cube mission has received Avoidance Applications Based On Machine Learning funding from the European Research Council (ERC) And Analytical Methods”, 8th European Conference under the European Union’s Horizon 2020 research and on Space Debris, ESA/ESOC, Darmstadt, Germany, innovation programme (grant agreement No 679086 - Virtual Conference, 20-23 April 2021. COMPAS 6. Horstman A., et al., “Enhancement of s/c

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